Turbine section of high bypass turbofan

ABSTRACT

A turbofan engine according to an example of the present disclosure includes, among other things, a fan including a circumferential array of fan blades, a compressor in fluid communication with the fan, the compressor including a low pressure compressor section and a high pressure compressor section, the low pressure compressor section including a low pressure compressor section inlet with a low pressure compressor section inlet annulus area, a fan duct including a fan duct annulus area outboard of the a low pressure compressor section inlet, a turbine in fluid communication with the combustor, the turbine having a high pressure turbine section and a low pressure turbine that drives the fan, a speed reduction mechanism coupled to the fan and rotatable by the low pressure turbine section to allow the low pressure turbine section to turn faster than the fan, wherein the low pressure turbine section includes a maximum gas path radius and the fan blades include a maximum radius, and a ratio of the maximum gas path radius to the maximum radius of the fan blades is between 0.50 and 0.55, or is greater than 0.55 and less than or equal to 0.65.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. patent application Ser. No.16/025,038, filed Jul. 2, 2018, which is a continuation-in-part of U.S.patent application Ser. No. 15/292,472, filed Oct. 13, 2016, which is acontinuation of U.S. patent application Ser. No. 14/793,785, filed Jul.8, 2015, which is a continuation-in-part of U.S. patent application Ser.No. 14/692,090, filed Apr. 21, 2015, which was a continuation of U.S.patent application Ser. No. 13/599,175, filed Aug. 30, 2012, which was acontinuation of U.S. patent application Ser. No. 13/475,252, now U.S.Pat. No. 8,844,265, issued Sep. 30, 2014, filed May 18, 2012, which wasa continuation-in-part of U.S. patent application Ser. No. 11/832,107,filed Aug. 1, 2007. U.S. patent application Ser. No. 13/475,252 claimedthe benefit of U.S. Patent Provisional Application No. 61/593,190, filedJan. 31, 2012, and U.S. Provisional Application No. 61/498,516, filedJun. 17, 2011.

BACKGROUND

The disclosure relates to turbofan engines. More particularly, thedisclosure relates to low pressure turbine sections of turbofan engineswhich power the fans via a speed reduction mechanism.

There has been a trend toward increasing bypass ratio in gas turbineengines. This is discussed further below. There has generally been acorrelation between certain characteristics of bypass and the diameterof the low pressure turbine section sections of turbofan engines.

SUMMARY

A turbofan engine according to an example of the present disclosureincludes a fan that has a circumferential array of fan blades. And acompressor in fluid communication with the fan. The compressor has afirst compressor section and a second compressor that has four stages.The second compressor has a greater number of stages than the firstcompressor. The second compressor section has a second compressorsection inlet with a second compressor section inlet annulus area. A fanduct has a fan duct annulus area outboard of the second compressorsection inlet. The ratio of the fan duct annulus area to the secondcompressor section inlet annulus area defines a bypass area ratio thatis greater than or equal to 8.0. A combustor is in fluid communicationwith the compressor. A shaft assembly has a first portion and a secondportion. A turbine is in fluid communication with the combustor. Theturbine has a two-stage first turbine section coupled to the firstportion of the shaft assembly to drive the first compressor section, anda four-stage second turbine section coupled to the second portion of theshaft assembly to drive the fan. Each of the second turbine section hasblades and vanes, and a second turbine airfoil count is defined as thenumerical count of all of the blades and vanes in the second turbinesection. An epicyclic transmission is coupled to the fan and rotatableby the second turbine section through the second portion of the shaftassembly to allow the second turbine to turn faster than the fan. Thesecond turbine airfoil count is below 1600. A ratio of the secondturbine airfoil count to the bypass area ratio is less than 150. Thesecond turbine section further includes a maximum gas path radius andthe fan blades include a maximum radius, and a ratio of the maximum gaspath radius to the maximum radius of the fan blades is less than 0.6.

A further embodiment of any of the foregoing embodiments includes a fancase and vanes, the fan case encircling the fan and supported by thevanes.

In a further embodiment of any of the foregoing embodiments, the fan isa single fan, and each fan blade includes a platform and an outboard endhaving a free tip.

In a further embodiment of any of the foregoing embodiments, theepicyclic transmission is a planetary gearbox has a speed reductionratio between 2:1 and 13:1 determined by the ratio of diameters withinthe planetary gearbox.

In a further embodiment of any of the foregoing embodiments, the ratioof the maximum gas path radius to the maximum radius of the fan bladesis greater than 0.55.

A further embodiment of any of the foregoing embodiments includes anengine aft mount location configured to support an engine mount when theengine is mounted and react at least a thrust load of the engine, and anengine forward mount location.

In a further embodiment of any of the foregoing embodiments, the engineforward mount location is axially proximate to the epicyclictransmission.

In a further embodiment of any of the foregoing embodiments, the engineforward mount location engages with an intermediate case.

In a further embodiment of any of the foregoing embodiments, the engineaft mount location engages with an engine thrust case.

In a further embodiment of any of the foregoing embodiments, the engineaft mount location is located between the second turbine section and thefirst turbine section.

In a further embodiment of any of the foregoing embodiments, the engineaft mount location is located between the second turbine section and thefirst turbine section.

In a further embodiment of any of the foregoing embodiments, the secondturbine section includes a plurality of blade stages interspersed with aplurality of vane stages, and each stage of the second turbine sectionincludes a disk with a circumferential array of blades. Each blade hasan airfoil extending from an inner diameter to an outer diameter. Theinner diameter is associated with a platform and the outer diameter isassociated with a shroud.

In a further embodiment of any of the foregoing embodiments, in at leastone stage the shroud is integral with the airfoil.

In a further embodiment of any of the foregoing embodiments, the shroudincludes outboard sealing ridges configured to seal with abradableseals.

In a further embodiment of any of the foregoing embodiments, theabradable seals include honeycomb.

A further embodiment of any of the foregoing embodiments includes a caseassociated with the second turbine section. The abradable seals arefixed to the case.

In a further embodiment of any of the foregoing embodiments, each stageof the second turbine section includes a disk, with a circumferentialarray of blades, each blade that has an airfoil extending from an innerdiameter to an outer diameter. The inner diameter is associated with aplatform and the outer diameter is unshrouded.

A further embodiment of any of the foregoing embodiments includes astationary blade outer air seal, and a rotational gap between a tip ofthe airfoil and the stationary blade outer air seal.

In a further embodiment of any of the foregoing embodiments, each of theplurality of vane stages includes a vane. Each vane has an airfoilextending from an inner diameter to an outer diameter. The innerdiameter is associated with a platform and the outer diameter isassociated with a shroud.

A further embodiment of any of the foregoing embodiments includes a caseassociated with the second turbine section. The shroud is fixed to thecase.

In a further embodiment of any of the foregoing embodiments, ahub-to-tip ratio (Ri:Ro) of the second turbine section is between 0.4and 0.5 measured at the maximum Ro axial location in the second turbinesection.

In a further embodiment of any of the foregoing embodiments, the bypassarea ratio is less than or equal to 20. The epicyclic transmission is aplanetary gearbox that has a speed reduction ratio between 2:1 and 13:1determined by the ratio of diameters within the planetary gearbox, andthe hub-to-tip ratio (Ri:Ro) is between 0.42-0.48.

In a further embodiment of any of the foregoing embodiments, the ratioof the maximum gas path radius to the maximum radius of the fan bladesis greater than 0.55, and the ratio of the second turbine airfoil countto the bypass area ratio is between 120 and 140.

A turbofan engine according to an example of the present disclosureincludes a fan that has a circumferential array of fan blades, and acompressor in fluid communication with the fan. The compressor has asecond compressor section and a first compressor section. The secondcompressor section has a second compressor section inlet with a secondcompressor section inlet annulus area. A fan duct has a fan duct annulusarea outboard of the second compressor section inlet. The ratio of thefan duct annulus area to the second compressor section inlet annulusarea defines a bypass area ratio that is greater than 6.0. A combustorin fluid communication with the compressor. A shaft assembly has a firstportion and a second portion. A turbine in fluid communication with thecombustor. The turbine has a first turbine section coupled to the firstportion of the shaft assembly to drive the first compressor section, anda second turbine section coupled to the second portion of the shaftassembly to drive the fan. Each of the second turbine section has bladesand vanes, and a second turbine airfoil count defined as the numericalcount of all of the blades and vanes in the second turbine section. Anepicyclic transmission is coupled to the fan and rotatable by the secondturbine section through the second portion of the shaft assembly toallow the second turbine to turn faster than the fan. The second turbineairfoil count is below 1600. A ratio of the second turbine airfoil countto the bypass area ratio is less than 150. The second turbine sectionfurther includes a maximum gas path radius and the fan blades include amaximum radius, and a ratio of the maximum gas path radius to themaximum radius of the fan blades is greater than or equal to 0.35. Ahub-to-tip ratio (Ri:Ro) of the second turbine section is between0.42-0.48 measured at the maximum Ro axial location in the secondturbine section.

In a further embodiment of any of the foregoing embodiments, the ratioof the maximum gas path radius to the maximum radius of the fan bladesis less than or equal to 0.6, the first turbine is a two-stage firstturbine and the second turbine is a four-stage second turbine.

In a further embodiment of any of the foregoing embodiments, the firstcompressor is a nine-stage first compressor.

In a further embodiment of any of the foregoing embodiments, the secondcompressor is a four-stage second compressor.

In a further embodiment of any of the foregoing embodiments, the bypassarea ratio is between 8.0 and 20, and the epicyclic transmission is aplanetary gearbox having a speed reduction ratio between 2:1 and 13:1determined by the ratio of diameters within the planetary gearbox.

In a further embodiment of any of the foregoing embodiments, the ratioof the maximum gas path radius to the maximum radius of the fan bladesis greater than 0.55, and the ratio of the second turbine airfoil countto the bypass area ratio is between 120 and 140.

A further embodiment of any of the foregoing embodiments includes anengine intermediate case, that has an engine forward mount locationproximate to the gearbox and configured to support an engine mount whenthe engine is mounted, and an engine thrust case that has an engine aftmount location configured to support an engine mount and react at leasta thrust load when the engine is mounted. The engine aft mount locationis located between the second turbine section and the first turbinesection.

A turbofan engine according to an example of the present disclosureincludes a fan that has a circumferential array of fan blades, and acompressor in fluid communication with the fan. The compressor has afirst compressor section and a second compressor that has four stages.The second compressor has a greater number of stages than the firstcompressor. The second compressor section has a second compressorsection inlet with a second compressor section inlet annulus area. A fanduct has a fan duct annulus area outboard of the second compressorsection inlet. The ratio of the fan duct annulus area to the secondcompressor section inlet annulus area defines a bypass area ratio thatis greater than or equal to 8.0. A combustor is in fluid communicationwith the compressor. A shaft assembly has a first portion and a secondportion. A turbine is in fluid communication with the combustor. Theturbine has a two-stage first turbine section coupled to the firstportion of the shaft assembly to drive the first compressor section, anda four-stage second turbine section coupled to the second portion of theshaft assembly to drive the fan. Each of the second turbine section hasblades and vanes, and a second turbine airfoil count defined as thenumerical count of all of the blades and vanes in the second turbinesection. An epicyclic transmission is coupled to the fan and rotatableby the second turbine section through the second portion of the shaftassembly to allow the second turbine to turn faster than the fan. Thesecond turbine airfoil count is below 1600. A ratio of the secondturbine airfoil count to the bypass area ratio is less than 150. Thesecond turbine section further includes a maximum gas path radius andthe fan blades that has a maximum radius, and a ratio of the maximum gaspath radius to the maximum radius of the fan blades is less than 0.6.

A further embodiment of any of the foregoing embodiments includes a fancase and vanes, the fan case encircling the fan and supported by thevanes.

In a further embodiment of any of the foregoing embodiments, the fan isa single fan, and each fan blade includes a platform and an outboard endhaving a free tip.

In a further embodiment of any of the foregoing embodiments, theepicyclic transmission is a planetary gearbox having a speed reductionratio between 2:1 and 13:1 determined by the ratio of diameters withinthe planetary gearbox.

In a further embodiment of any of the foregoing embodiments, the ratioof the maximum gas path radius to the maximum radius of the fan bladesis greater than 0.55.

A further embodiment of any of the foregoing embodiments includes anengine aft mount location configured to support an engine mount when theengine is mounted and react at least a thrust load of the engine, and anengine forward mount location.

In a further embodiment of any of the foregoing embodiments, the engineforward mount location is axially proximate to the epicyclictransmission.

In a further embodiment of any of the foregoing embodiments, the engineforward mount location engages with an intermediate case.

In a further embodiment of any of the foregoing embodiments, the engineaft mount location engages with an engine thrust case.

In a further embodiment of any of the foregoing embodiments, the engineaft mount location is located between the second turbine section and thefirst turbine section.

In a further embodiment of any of the foregoing embodiments, the engineaft mount location is located between the second turbine section and thefirst turbine section.

In a further embodiment of any of the foregoing embodiments, the secondturbine section includes a plurality of blade stages interspersed with aplurality of vane stages, and each stage of the second turbine sectionincludes a disk with a circumferential array of blades. Each blade hasan airfoil extending from an inner diameter to an outer diameter. Theinner diameter is associated with a platform and the outer diameter isassociated with a shroud.

In a further embodiment of any of the foregoing embodiments, in at leastone stage the shroud is integral with the airfoil.

In a further embodiment of any of the foregoing embodiments, the shroudincludes outboard sealing ridges configured to seal with abradableseals.

In a further embodiment of any of the foregoing embodiments, theabradable seals include honeycomb.

A further embodiment of any of the foregoing embodiments includes a caseassociated with the second turbine section. The abradable seals arefixed to the case.

In a further embodiment of any of the foregoing embodiments, each stageof the second turbine section includes a disk, with a circumferentialarray of blades, each blade that has an airfoil extending from an innerdiameter to an outer diameter. The inner diameter is associated with aplatform and the outer diameter is unshrouded.

A further embodiment of any of the foregoing embodiments includes astationary blade outer air seal, and a rotational gap between a tip ofthe airfoil and the stationary blade outer air seal.

In a further embodiment of any of the foregoing embodiments, each of theplurality of vane stages includes a vane, each vane has an airfoilextending from an inner diameter to an outer diameter. The innerdiameter is associated with a platform and the outer diameter isassociated with a shroud.

A further embodiment of any of the foregoing embodiments includes a caseassociated with the second turbine section, wherein the shroud is fixedto the case.

In a further embodiment of any of the foregoing embodiments, ahub-to-tip ratio (Ri:Ro) of the second turbine section is between 0.4and 0.5 measured at the maximum Ro axial location in the second turbinesection.

In a further embodiment of any of the foregoing embodiments, the bypassarea ratio is less than or equal to 20, the epicyclic transmission is aplanetary gearbox having a speed reduction ratio between 2:1 and 13:1determined by the ratio of diameters within the planetary gearbox, andthe hub-to-tip ratio (Ri:Ro) is between 0.42-0.48.

In a further embodiment of any of the foregoing embodiments, the ratioof the maximum gas path radius to the maximum radius of the fan bladesis greater than 0.55, and the ratio of the second turbine airfoil countto the bypass area ratio is between 120 and 140.

A turbofan engine according to an example of the present disclosureincludes a fan that has a circumferential array of fan blades, and acompressor in fluid communication with the fan. The compressor has asecond compressor section and a first compressor section. The secondcompressor section has a second compressor section inlet with a secondcompressor section inlet annulus area, and a fan duct that has a fanduct annulus area outboard of the second compressor section inlet. Theratio of the fan duct annulus area to the second compressor sectioninlet annulus area defines a bypass area ratio that is greater than 6.0.A combustor in fluid communication with the compressor. A shaft assemblyhas a first portion and a second portion. A turbine in fluidcommunication with the combustor. The turbine has a first turbinesection coupled to the first portion of the shaft assembly to drive thefirst compressor section, and a second turbine section coupled to thesecond portion of the shaft assembly to drive the fan. Each of thesecond turbine section includes blades and vanes, and a second turbineairfoil count defined as the numerical count of all of the blades andvanes in the second turbine section. An epicyclic transmission iscoupled to the fan and rotatable by the second turbine section throughthe second portion of the shaft assembly to allow the second turbine toturn faster than the fan. The second turbine airfoil count is below1600. A ratio of the second turbine airfoil count to the bypass arearatio is less than 150. The second turbine section further includes amaximum gas path radius and the fan blades include a maximum radius, anda ratio of the maximum gas path radius to the maximum radius of the fanblades is greater than or equal to 0.35. A hub-to-tip ratio (Ri:Ro) ofthe second turbine section is between 0.42-0.48 measured at the maximumRo axial location in the second turbine section.

In a further embodiment of any of the foregoing embodiments, the ratioof the maximum gas path radius to the maximum radius of the fan bladesis less than or equal to 0.6, the first turbine is a two-stage firstturbine and the second turbine is a four-stage second turbine.

In a further embodiment of any of the foregoing embodiments, the firstcompressor is a nine-stage first compressor.

In a further embodiment of any of the foregoing embodiments, the secondcompressor is a four-stage second compressor.

In a further embodiment of any of the foregoing embodiments, the bypassarea ratio is between 8.0 and 20, and the epicyclic transmission is aplanetary gearbox having a speed reduction ratio between 2:1 and 13:1determined by the ratio of diameters within the planetary gearbox.

In a further embodiment of any of the foregoing embodiments, the ratioof the maximum gas path radius to the maximum radius of the fan bladesis greater than 0.55, and the ratio of the second turbine airfoil countto the bypass area ratio is between 120 and 140.

A further embodiment of any of the foregoing embodiments includes anengine intermediate case that has an engine forward mount locationproximate to the gearbox and configured to support an engine mount whenthe engine is mounted, and an engine thrust case that has an engine aftmount location configured to support an engine mount and react at leasta thrust load when the engine is mounted. The engine aft mount locationis located between the second turbine section and the first turbinesection.

One aspect of the disclosure involves a turbofan engine having an enginecase and a gaspath through the engine case. A fan has a circumferentialarray of fan blades. The engine further has a compressor in fluidcommunication with the fan, a combustor in fluid communication with thecompressor, a turbine in fluid communication with the combustor, whereinthe turbine includes a low pressure turbine section having 3 to 6 bladestages. A speed reduction mechanism couples the low pressure turbinesection to the fan. A bypass area ratio is greater than about 6.0. Aratio of the total number of airfoils in the low pressure turbinesection divided by the bypass area ratio is less than about 170, saidlow pressure turbine section airfoil count being the total number ofblade airfoils and vane airfoils of the low pressure turbine section.

In additional or alternative embodiments of any of the foregoingembodiments, the bypass area ratio may be greater than about 8.0 or maybe between about 8.0 and about 20.0.

In additional or alternative embodiments of any of the foregoingembodiments, a fan case may encircle the fan blades radially outboard ofthe engine case.

In additional or alternative embodiments of any of the foregoingembodiments, the compressor may comprise a low pressure compressorsection and a high pressure compressor section.

In additional or alternative embodiments of any of the foregoingembodiments, the blades of the low pressure compressor section and lowpressure turbine section may share a low shaft.

In additional or alternative embodiments of any of the foregoingembodiments, the high pressure compressor section and a high pressureturbine section of the turbine may share a high shaft.

In additional or alternative embodiments of any of the foregoingembodiments, there are no additional compressor or turbine sections.

In additional or alternative embodiments of any of the foregoingembodiments, the speed reduction mechanism may comprise an epicyclictransmission coupling the low speed shaft to a fan shaft to drive thefan with a speed reduction.

In additional or alternative embodiments of any of the foregoingembodiments, the low pressure turbine section may have an exemplary 2 to6 blade stages or 2 to 3 blade stages.

In additional or alternative embodiments of any of the foregoingembodiments, a hub-to-tip ratio (Ri:Ro) of the low pressure turbinesection may be between about 0.4 and about 0.5 measured at the maximumRo axial location in the low pressure turbine section.

In additional or alternative embodiments of any of the foregoingembodiments, a ratio of maximum gaspath radius along the low pressureturbine section to maximum radius of the fan may be less than about0.55, or less than about 0.50, or between about 0.35 and about 0.50.

In additional or alternative embodiments of any of the foregoingembodiments, the ratio of low pressure turbine section airfoil count tobypass area ratio may be between about 10 and about 150.

In additional or alternative embodiments of any of the foregoingembodiments, the airfoil count of the low pressure turbine section maybe below about 1600.

In additional or alternative embodiments of any of the foregoingembodiments, the engine may be in combination with a mountingarrangement (e.g., of an engine pylon) wherein an aft mount reacts atleast a thrust load.

The details of one or more embodiments are set forth in the accompanyingdrawings and the description below. Other features, objects, andadvantages will be apparent from the description and drawings, and fromthe claims.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is an axial sectional view of a turbofan engine.

FIG. 2 is an axial sectional view of a low pressure turbine section ofthe engine of FIG. 1.

FIG. 3 is transverse sectional view of transmission of the engine ofFIG. 1.

FIG. 4 shows another embodiment.

FIG. 5 shows yet another embodiment.

Like reference numbers and designations in the various drawings indicatelike elements.

DETAILED DESCRIPTION

FIG. 1 shows a turbofan engine 20 having a main housing (engine case) 22containing a rotor shaft assembly 23. Via high 24 and low 25 shaftportions of the shaft assembly 23, a high pressure turbine section (gasgenerating turbine) 26 and a low pressure turbine section 27respectively drive a high pressure compressor section 28 and a lowpressure compressor section 30. As used herein, the high pressureturbine section experiences higher pressures that the low pressureturbine section. A low pressure turbine section is a section that powersa fan 42. Although a two-spool (plus fan) engine is shown, one of manyalternative variations involves a three-spool (plus fan) engine whereinan intermediate spool comprises an intermediate pressure compressorbetween the low fan and high pressure compressor section and anintermediate pressure turbine between the high pressure turbine sectionand low pressure turbine section.

The engine extends along a longitudinal axis 500 from a fore end to anaft end. Adjacent the fore end, a shroud (fan case) 40 encircles the fan42 and is supported by vanes 44. An aerodynamic nacelle around the fancase is shown and an aerodynamic nacelle 45 around the engine case isshown.

The low shaft portion 25 of the rotor shaft assembly 23 drives the fan42 through a speed reduction mechanism 46. An exemplary speed reductionmechanism is an epicyclic transmission, namely a star or planetary gearsystem. As is discussed further below, an inlet airflow 520 entering thenacelle is divided into a portion 522 passing along a core flowpath 524and a bypass portion 526 passing along a bypass flowpath 528. With theexception of diversions such as cooling air, etc., flow along the coreflowpath sequentially passes through the low pressure compressorsection, high pressure compressor section, a combustor 48, the highpressure turbine section, and the low pressure turbine section beforeexiting from an outlet 530.

FIG. 3 schematically shows details of the transmission 46. A forward endof the low shaft 25 is coupled to a sun gear 52 (or other high speedinput to the speed reduction mechanism). The externally-toothed sun gear52 is encircled by a number of externally-toothed star gears 56 and aninternally-toothed ring gear 54. The exemplary ring gear is coupled tothe fan to rotate with the fan as a unit.

The star gears 56 are positioned between and enmeshed with the sun gearand ring gear. A cage or star carrier assembly 60 carries the star gearsvia associated journals 62. The exemplary star carrier is substantiallyirrotatably mounted relative via fingers 404 to the case 22.

Another transmission/gearbox combination has the star carrier connectedto the fan and the ring is fixed to the fixed structure (case) ispossible and such is commonly referred to as a planetary gearbox.

The speed reduction ratio is determined by the ratio of diameters withinthe gearbox. An exemplary reduction is between about 2:1 and about 13:1.

The exemplary fan (FIG. 1) comprises a circumferential array of blades70. Each blade comprises an airfoil 72 having a leading edge 74 and atrailing edge 76 and extending from an inboard end 78 at a platform toan outboard end 80 (i.e., a free tip). The outboard end 80 is in closefacing proximity to a rub strip 82 along an interior surface 84 of thenacelle and fan case.

To mount the engine to the aircraft wing 92, a pylon 94 is mounted tothe fan case and/or to the other engine cases. The exemplary pylon 94may be as disclosed in U.S. patent application Ser. No. 11/832,107(US2009/0056343A1). The pylon comprises a forward mount 100 and anaft/rear mount 102. The forward mount may engage the engine intermediatecase (IMC) and the aft mount may engage the engine thrust case. The aftmount reacts at least a thrust load of the engine.

To reduce aircraft fuel burn with turbofans, it is desirable to producea low pressure turbine with the highest efficiency and lowest weightpossible. Further, there are considerations of small size (especiallyradial size) that benefit the aerodynamic shape of the engine cowlingand allow room for packaging engine subsystems.

FIG. 2 shows the low pressure turbine section 27 as comprising anexemplary three blade stages 200, 202, 204. An exemplary blade stagecount is 2-6, more narrowly, 2-4, or 2-3, 3-5, or 3-4. Interspersedbetween the blade stages are vane stages 206 and 208. Each exemplaryblade stage comprises a disk 210, 212, and 214, respectively. Acircumferential array of blades extends from peripheries of each of thedisks. Each exemplary blade comprises an airfoil 220 extending from aninner diameter (ID) platform 222 to an outer diameter (OD) shroud 224(shown integral with the airfoil).

An alternative may be an unshrouded blade with a rotational gap betweenthe tip of the blade and a stationary blade outer air seal (BOAS). Eachexemplary shroud 224 has outboard sealing ridges which seal withabradable seals (e.g., honeycomb) fixed to the case. The exemplary vanesin stages 206 and 208 include airfoils 230 extending from ID platforms232 to OD shrouds 234. The exemplary OD shrouds 234 are directly mountedto the case. The exemplary platforms 232 carry seals for sealing withinter-disk knife edges protruding outwardly from inter-disk spacerswhich may be separate from the adjacent disks or unitarily formed withone of the adjacent disks.

Each exemplary disk 210, 212, 214 comprises an enlarged central annularprotuberance or “bore” 240, 242, 244 and a thinner radial web 246, 248,250 extending radially outboard from the bore. The bore impartsstructural strength allowing the disk to withstand centrifugal loadingwhich the disk would otherwise be unable to withstand.

A turbofan engine is characterized by its bypass ratio (mass flow ratioof air bypassing the core to air passing through the core) and thegeometric bypass area ratio (ratio of fan duct annulus areaoutside/outboard of the low pressure compressor section inlet (i.e., atlocation 260 in FIG. 1) to low pressure compressor section inlet annulusarea (i.e., at location 262 in FIG. 2). High bypass engines typicallyhave bypass area ratio of at least four. There has been a correlationbetween increased bypass area ratio and increased low pressure turbinesection radius and low pressure turbine section airfoil count. As isdiscussed below, this correlation may be broken by having an engine withrelatively high bypass area ratio and relatively low turbine size.

In some embodiments, the engine 20 bypass ratio is greater than or equalto about six (6), with an example embodiment being greater than or equalto about ten (10), the speed reduction mechanism 46 defines a gearreduction ratio of greater than about 2.3 and the low pressure turbinesection 27 has a pressure ratio that is greater than about five. In onefurther non-limiting embodiment, the low pressure turbine section 27 hasa pressure ratio that is greater than about five and less than aboutten. In one disclosed embodiment, the engine 20 bypass ratio is greaterthan about ten (10:1), the fan diameter is significantly larger thanthat of the low pressure compressor section 30, and the low pressureturbine section 27 has a pressure ratio that is greater than about five5:1. Low pressure turbine section 27 pressure ratio is pressure measuredprior to inlet of low pressure turbine section 27 as related to thepressure at the outlet of the low pressure turbine section 27 prior toan exhaust nozzle. The low pressure turbine section 27 can have apressure ratio that is less than or equal to about 20.0, such as betweenabout 10.0 and about 15.0. In another embodiment, the engine 20 has abypass ratio less than or equal to about 25.0, such as between about15.0 and about 20.0, or between about 15.0 and 18.0. The gear reductionratio can be less than about 5.0, or less than about 4.0, for example,or between about 4.0 and 5.0.

The fan 42 is designed for a particular flight condition—typicallycruise at about 0.8 Mach and about 35,000 feet. The flight condition of0.8 Mach and 35,000 ft, with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.50, or more narrowly less than about1.45, or between about 1.3 and 1.45, or between about 1.30 and 1.38. Inembodiments, the low fan pressure ratio is greater than or equal toabout 1.1 or about 1.2. “Low corrected fan tip speed” is the actual fantip speed in ft/sec divided by an industry standard temperaturecorrection of [(Tram ° R)/(518.7° R)]^(0.5). The “Low corrected fan tipspeed” as disclosed herein according to one non-limiting embodiment isless than about 1150 ft/second. Further, low corrected fan tip speedaccording to one non-limiting embodiment is greater than about 1000ft/second.

By employing a speed reduction mechanism (e.g., a transmission) to allowthe low pressure turbine section to turn very fast relative to the fanand by employing low pressure turbine section design features for highspeed, it is possible to create a compact turbine module (e.g., whileproducing the same amount of thrust and increasing bypass area ratio).The exemplary transmission is an epicyclic transmission. Alternativetransmissions include composite belt transmissions, metal chain belttransmissions, fluidic transmissions, and electric means (e.g., amotor/generator set where the turbine turns a generator providingelectricity to an electric motor which drives the fan).

Compactness of the turbine is characterized in several ways. Along thecompressor and turbine sections, the core gaspath extends from aninboard boundary (e.g., at blade hubs or outboard surfaces of platformsof associated blades and vanes) to an outboard boundary (e.g., at bladetips and inboard surfaces of blade outer air seals for unshrouded bladetips and at inboard surfaces of OD shrouds of shrouded blade tips and atinboard surfaces of OD shrouds of the vanes). These boundaries may becharacterized by radii R_(I) and Ro, respectively, which vary along thelength of the engine.

For low pressure turbine radial compactness, there may be a relativelyhigh ratio of radial span (R_(O)-R_(I)) to radius (R_(O) or R_(I)).Radial compactness may also be expressed in the hub-to-tip ratio(R_(I):R_(O)). These may be measured at the maximum Ro location in thelow pressure turbine section. The exemplary compact low pressure turbinesection has a hub-to-tip ratio close to about 0.5 (e.g., about 0.4-0.5or about 0.42-0.48, with an exemplary about 0.46).

Another characteristic of low pressure turbine radial compactness isrelative to the fan size. An exemplary fan size measurement is themaximum tip radius R_(Tmax) of the fan blades. An exemplary ratio is themaximum Ro along the low pressure turbine section to R_(Tmax) of the fanblades. Exemplary values for this ratio are less than or equal to about0.65, or more narrowly, less than or equal to about 0.6, above about0.55, less than or equal to about 0.55 (e.g., about 0.35-0.55), lessthan about 0.50, or about 0.35-0.50.

To achieve compactness the designer may balance multiple physicalphenomena to arrive at a system solution as defined by the low pressureturbine hub-to-tip ratio, the fan maximum tip radius to low pressureturbine maximum Ro ratio, the bypass area ratio, and the bypass arearatio to low pressure turbine airfoil count ratio. These concernsinclude, but are not limited to: a) aerodynamics within the low pressureturbine, b) low pressure turbine blade structural design, c) lowpressure turbine disk structural design, and d) the shaft connecting thelow pressure turbine to the low pressure compressor and speed reductiondevice between the low pressure compressor and fan. These physicalphenomena may be balanced in order to achieve desirable performance,weight, and cost characteristics.

The addition of a speed reduction device between the fan and the lowpressure compressor creates a larger design space because the speed ofthe low pressure turbine is decoupled from the fan. This design spaceprovides great design variables and new constraints that limitfeasibility of a design with respect to physical phenomena. For examplethe designer can independently change the speed and flow area of the lowpressure turbine to achieve optimal aerodynamic parameters defined byflow coefficient (axial flow velocity/wheel speed) and work coefficient(wheel speed/square root of work). However, this introduces structuralconstraints with respect blade stresses, disk size, material selection,etc.

In some examples, the designer can choose to make low pressure turbinesection disk bores much thicker relative to prior art turbine bores andthe bores may be at a much smaller radius RB. This increases the amountof mass at less than a “self sustaining radius”. Another means is tochoose disk materials of greater strength than prior art such as the useof wrought powdered metal disks to allow for extremely high centrifugalblade pulls associated with the compactness.

Another variable in achieving compactness is to increase the structuralparameter AN² which is the annulus area of the exit of the low pressureturbine divided by the low pressure turbine rpm squared at its redlineor maximum speed. Relative to prior art turbines, which are greatlyconstrained by fan blade tip mach number, a very wide range of AN²values can be selected and optimized while accommodating suchconstraints as cost or a countering, unfavorable trend in low pressureturbine section shaft dynamics. In selecting the turbine speed (andthereby selecting the transmission speed ratio, one has to be mindfulthat at too high a gear ratio the low pressure turbine section shaft(low shaft) will become dynamically unstable.

The higher the design speed, the higher the gear ratio will be and themore massive the disks will become and the stronger the low pressureturbine section disk and blade material will have to be. All of theseparameters can be varied simultaneously to change the weight of theturbine, its efficiency, its manufacturing cost, the degree ofdifficulty in packaging the low pressure turbine section in the corecowling and its durability. This is distinguished from a prior artdirect drive configuration, where the high bypass area ratio can only beachieved by a large low pressure turbine section radius. Because thatradius is so very large and, although the same variables (airfoilturning, disk size, blade materials, disk shape and materials, etc.) aretheoretically available, as a practical matter economics and engine fuelburn considerations severely limit the designer's choice in theseparameters.

Another characteristic of low pressure turbine section size is airfoilcount (numerical count of all of the blades and vanes in the lowpressure turbine). Airfoil metal angles can be selected such thatairfoil count is low or extremely low relative to a direct driveturbine. In known prior art engines having bypass area ratio above 6.0(e.g. 8.0-20), low pressure turbine sections involve ratios of airfoilcount to bypass area ratio above 190.

With the full range of selection of parameters discussed aboveincluding, disk bore thickness, disk material, hub to tip ratio, andR_(O)/R_(Tmax), the ratio of airfoil count to bypass area ratio may bebelow about 170 to as low as 10 (e.g., equal to or below about 150 or anexemplary about 10-170, more narrowly about 10-150, or above about 150by implication). In some embodiments, the ratio of airfoil count tobypass area ratio is greater than or equal to about 100, such as betweenabout 120 and 140, and the low pressure turbine section 27 has betweenthree and four stages. In other embodiments, the ratio of airfoil countto bypass area ratio is less than 100, such as between about 15 and 80,and the low pressure turbine section 27 has between three and fourstages. Further, in such embodiments the airfoil count may be belowabout 1700, or below about 1600 or below about 1000, such as about300-800 airfoils, or more narrowly between about 350-750 airfoils.

FIG. 4 shows an embodiment 600, wherein there is a fan drive turbine 608driving a shaft 606 to in turn drive a fan rotor 602. A gear reduction604 may be positioned between the fan drive turbine 608 and the fanrotor 602. This gear reduction 604 may be structured and operate likethe gear reduction disclosed above. A compressor rotor 610 is driven byan intermediate pressure turbine 612, and a second stage compressorrotor 614 is driven by a turbine rotor 216. A combustion section 618 ispositioned intermediate the compressor rotor 614 and the turbine section616.

FIG. 5 shows yet another embodiment 700 wherein a fan rotor 702 and afirst stage compressor 704 rotate at a common speed. The gear reduction706 (which may be structured as disclosed above) is intermediate thecompressor rotor 704 and a shaft 708 which is driven by a low pressureturbine section.

One or more embodiments have been described. Nevertheless, it will beunderstood that various modifications may be made. For example, whenreengineering from a baseline engine configuration, details of thebaseline may influence details of any particular implementation.Accordingly, other embodiments are within the scope of the followingclaims.

What is claimed is:
 1. A turbofan engine comprising: a fan including acircumferential array of fan blades, a fan case and vanes, the fan caseencircling the fan and supported by the vanes; a compressor in fluidcommunication with the fan, the compressor including a low pressurecompressor section and a high pressure compressor section, the highpressure compressor section including a greater number of stages thanthe low pressure compressor section, and the low pressure compressorsection including a low pressure compressor section inlet with a lowpressure compressor section inlet annulus area; a fan duct including afan duct annulus area outboard of the low pressure compressor sectioninlet, wherein the fan duct annulus area and the low pressure compressorsection inlet annulus area are established at a splitter that bounds thefan duct and the low pressure compressor section inlet, and wherein theratio of the fan duct annulus area to the low pressure compressorsection inlet annulus area defines a bypass area ratio that is greaterthan or equal to 8.0; a combustor in fluid communication with thecompressor; a shaft assembly having a first portion and a secondportion; a turbine in fluid communication with the combustor, theturbine having a high pressure turbine section coupled to the firstportion of the shaft assembly to drive the high pressure compressorsection, and a low pressure turbine section coupled to the secondportion of the shaft assembly to drive the fan, the high pressureturbine section including two stages, the low pressure compressorsection having a greater number of stages than the high pressure turbinesection, the low pressure turbine section including blades and vanes,and a low pressure turbine airfoil count defined as the numerical countof all of the blades and vanes in the low pressure turbine section; anda speed reduction mechanism coupled to the fan and rotatable by the lowpressure turbine section through the second portion of the shaftassembly to allow the low pressure turbine section to turn faster thanthe fan; wherein the low pressure turbine airfoil count is below 1600;wherein a ratio of the low pressure turbine airfoil count to the bypassarea ratio is less than 150; and wherein the low pressure turbinesection further includes a maximum gas path radius and the fan bladesinclude a maximum radius, and a ratio of the maximum gas path radius tothe maximum radius of the fan blades is greater than 0.55 and less thanor equal to 0.65.
 2. The turbofan engine as recited in claim 1, whereinthe speed reduction mechanism is an epicyclic transmission.
 3. Theturbofan engine as recited in claim 2, wherein the epicyclictransmission is axially between the splitter and a forwardmost row ofblades of the low pressure compressor section relative to a longitudinalaxis of the engine, and the fan is a single-stage fan.
 4. The turbofanengine as recited in claim 3, wherein a hub-to-tip ratio (Ri:Ro) of thelow pressure turbine section is between 0.42-0.48 measured at themaximum Ro axial location in the low pressure turbine section.
 5. Theturbofan engine as recited in claim 4, wherein the low pressure turbinesection is a three-stage or four-stage turbine.
 6. The turbofan engineas recited in claim 5, wherein the low pressure turbine airfoil count isbetween 300 and 800 airfoils.
 7. The turbofan engine as recited in claim3, wherein a hub-to-tip ratio (Ri:Ro) of the low pressure turbinesection is between 0.4-0.5 measured at the maximum Ro axial location inthe low pressure turbine section.
 8. The turbofan engine as recited inclaim 7, wherein the low pressure turbine section is a three-stage orfour-stage turbine.
 9. The turbofan engine as recited in claim 8,wherein the epicyclic transmission is a star gear system.
 10. Theturbofan engine as recited in claim 9, wherein the hub-to-tip ratio(Ri:Ro) is between 0.42-0.48.
 11. The turbofan engine as recited inclaim 10, further comprising a fan pressure ratio of less than 1.45measured across the fan blade alone at cruise at 0.8 Mach and 35,000 ft,wherein the ratio of the maximum gas path radius to the maximum radiusof the fan blades is less than 0.6.
 12. The turbofan engine as recitedin claim 11, wherein the low pressure turbine airfoil count is below1000.
 13. The turbofan engine as recited in claim 12, wherein the lowpressure compressor section and the low pressure turbine section includean equal number of stages.
 14. The turbofan engine as recited in claim12, wherein the low pressure compressor section has a greater number ofstages than the low pressure turbine section.
 15. The turbofan engine asrecited in claim 14, wherein the low pressure compressor section hasfour stages.
 16. The turbofan engine as recited in claim 8, wherein theepicyclic transmission is a planetary gear system.
 17. The turbofanengine as recited in claim 16, wherein the hub-to-tip ratio (Ri:Ro) isbetween 0.42-0.48.
 18. The turbofan engine as recited in claim 17,further comprising a fan pressure ratio of less than 1.45 measuredacross the fan blade alone at cruise at 0.8 Mach and 35,000 ft, whereinthe ratio of the maximum gas path radius to the maximum radius of thefan blades is less than 0.6.
 19. The turbofan engine as recited in claim18, wherein the low pressure turbine airfoil count is below
 1000. 20.The turbofan engine as recited in claim 19, wherein the low pressurecompressor section and the low pressure turbine section include an equalnumber of stages.
 21. The turbofan engine as recited in claim 19,wherein the low pressure compressor section has a greater number ofstages than the low pressure turbine section.
 22. The turbofan engine asrecited in claim 21, wherein the low pressure compressor section hasfour stages.
 23. A turbofan engine comprising: a fan including acircumferential array of fan blades, a fan case and vanes, the fan caseencircling the fan and supported by the vanes; a compressor in fluidcommunication with the fan, the compressor including a low pressurecompressor section and a high pressure compressor section, the lowpressure compressor section including a low pressure compressor sectioninlet with a low pressure compressor section inlet annulus area; a fanduct including a fan duct annulus area outboard of the low pressurecompressor section inlet, wherein the fan duct annulus area and the lowpressure compressor section inlet annulus area are established at asplitter that bounds the fan duct and the low pressure compressorsection inlet, and wherein the ratio of the fan duct annulus area to thelow pressure compressor section inlet annulus area defines a bypass arearatio that is greater than 8.0; a combustor in fluid communication withthe compressor; a shaft assembly having a first portion and a secondportion; a turbine in fluid communication with the combustor, theturbine having a high pressure turbine section coupled to the firstportion of the shaft assembly to drive the high pressure compressorsection, and a low pressure turbine section coupled to the secondportion of the shaft assembly to drive the fan, the low pressure turbinesection including blades and vanes, and a low pressure turbine airfoilcount defined as the numerical count of all of the blades and vanes inthe low pressure turbine section; and a speed reduction mechanismcoupled to the fan and rotatable by the low pressure turbine sectionthrough the second portion of the shaft assembly to allow the lowpressure turbine section to turn faster than the fan; wherein the lowpressure turbine airfoil count is below 1600; wherein a ratio of the lowpressure turbine airfoil count to the bypass area ratio is less than150; wherein the low pressure turbine section further includes a maximumgas path radius and the fan blades include a maximum radius, and a ratioof the maximum gas path radius to the maximum radius of the fan bladesis between 0.50 and 0.55; and wherein a hub-to-tip ratio (Ri:Ro) of thelow pressure turbine section is between 0.42-0.48 measured at themaximum Ro axial location in the low pressure turbine section.
 24. Theturbofan engine as recited in claim 23, wherein the high pressureturbine section includes two stages, and the low pressure compressorsection has a greater number of stages than the high pressure turbinesection.
 25. The turbofan engine as recited in claim 24, wherein thespeed reduction mechanism is an epicyclic transmission.
 26. The turbofanengine as recited in claim 25, wherein the epicyclic transmission isaxially between the splitter and a forwardmost row of blades of the lowpressure compressor section relative to a longitudinal axis of theengine, and the fan is a single-stage fan.
 27. The turbofan engine asrecited in claim 26, wherein the low pressure turbine section drives thelow pressure compressor section and an input of the epicyclictransmission through the second portion of the shaft assembly.
 28. Theturbofan engine as recited in claim 27, wherein the low pressure turbinesection is a four-stage turbine.
 29. The turbofan engine as recited inclaim 27, wherein the low pressure turbine section is a three-stageturbine.
 30. The turbofan engine as recited in claim 29, furthercomprising a fan pressure ratio of less than 1.45 measured across thefan blade alone at cruise at 0.8 Mach and 35,000 ft, wherein the lowpressure turbine section includes an inlet, an outlet, and a pressureratio of greater than 5, the pressure ratio being pressure measuredprior to the inlet as related to pressure at the outlet prior to anyexhaust nozzle.